The present invention relates generally to gas turbine engines, and, more specifically, to turbine rotor blades therein.
A typical gas turbine engine includes a compressor for pressuring air which is mixed with fuel in a combustor and ignited for generating hot combustion gases which flow downstream through a turbine which extracts energy therefrom for powering the compressor. The turbine includes a plurality of circumferentially adjoining rotor blades extending radially outwardly from the perimeter of a supporting disk.
A typical turbine blade includes an airfoil having a generally concave pressure side and an opposite, generally convex suction side extending axially between opposite leading and trailing edges which extend radially from a root to tip of the airfoil. The blade also includes a platform integrally joined to the root of the airfoil which defines a radially inner flowpath boundary for the combustion gases. Extending radially below the platform is an integral dovetail which slidingly engages a complementary dovetail slot extending axially through the rotor disk for retention of the blade during operation.
The turbine blades and rotor disk require precise dimensions for maximizing aerodynamic efficiency of the turbine and limiting stress during operation from centrifugal force, pressure loads, and thermal gradients;. However, since the turbine blades and disk are individually manufactured, they are subject to statistical variation in their dimensions, including statistical variation in stack-up tolerances when the blades are assembled into the disk.
For example, the individual blade platforms collectively define the radially inner flowpath for the combustion gases channeled over the turbine airfoils. The radial location of the outer surface of the platforms from the axial centerline axis of the turbine varies randomly from platform to platform around the circumference of the disk. Accordingly, some platforms are radially higher than adjacent platforms and some are radially lower, and in both situations effect differential steps therebetween along the circumferential side edges of the platforms. As the combustion gases engage the steps, aerodynamic efficiency may be adversely affected, and the protruding steps are locally heated by the hot combustion gases. This local heating can adversely affect the useful life of the blades and is undesirable, especially for turbines operated at ever increasing combustion gas temperatures.
The adverse affect of the steps is ameliorated by providing a chamfer which extends along both circumferential side edges of the individual platforms. The chamfers provide relatively smooth transitions from platform to platform notwithstanding the small differences in radial position of the adjacent platforms. Such chamfered turbine blade platforms have enjoyed many years of successful commercial use in this country. However, the chamfers require additional manufacturing steps and cost and introduce yet another feature which must accurately controlled during manufacture.
More specifically, modern turbine blades are relatively complex and expensive to manufacture since they are typically made of high temperature, high strength superalloy materials. The blades are typically hollow and include various internal cooling features therein along which a portion of the pressurized air bled from the compressor is channeled, and typically discharged through the airfoil through various film cooling and other holes drilled through any one of the sides, leading and trailing edges, and tip thereof.
Turbine blades are typically cast to near final shape and dimension in a conventional lost wax method. The process starts with a master mold or wax die in which is initially cast a wax form of the entire blade. The internal cooling features of the blade are separately formed in a corresponding core. The core and wax blade are then placed in a suitable mold, and the molten metal displaces the wax around the core and solidifies to form the cast blade.
The cast metal blade then undergoes additional manufacturing steps to obtain the final or finished dimensions thereof, and various holes may then be drilled through the airfoil as required. Since the blades are disposed in a row around the perimeter of the rotor disk, the circumferential width of the individual platforms requires precise dimensions and tolerances to prevent excessively large or narrow gaps therebetween when assembled.
The platform side edges are typically machined to final dimension using a precision grinder. The edge chamfers are then separately formed by using another suitable grinder, such as a pencil grinder, for manually blunting the finished platform side edges to form the chamfers thereat.
This chamfering requires suitable care, and attendant additional cost, to prepare the platforms in final dimension. And, it is subject to its own manufacturing variations. For example, the chamfers should be uniform in extent along the entire side edges of the platforms for accommodating the statistical differences in radial position thereof from platform to platform.
And, since the trailing edge of the individual airfoils is disposed closely adjacent to one of the side edges, the chamfering in this region must be carefully effected to prevent damage to the trailing edge. The trailing edge is subject to high temperature during operation and high stress, and damage thereof where it adjoins the platform edge may require scrapping of the entire blade, with a corresponding waste of manufacturing effort and expense.
Accordingly, it is desired to provide an improved method of making a gas turbine blade with platform chamfers.